Rocket engine nozzle

ABSTRACT

A rocket engine system comprising a single nozzle having a curved profile in axial section. In order to enable a control of the flow separation occurring within the nozzle outlet, the outlet portion in transverse section has a radius, the length of which varies such that the transverse section transitions smoothly from a circular nozzle throat to a polygonal outlet section.

The present invention refers to a rocket engine nozzle with an outletportion or thrust chamber having a curved profile in axial section.

During start-up and stop transients in sea-level rocket enginessignificant dynamic loads usually occur. These loads are generallyattributed to the disordered flow characteristics of the flow duringflow separation.

The outlet portion of nozzles for liquid propellant rocket engines oftenoperate at conditions where the main jet exhausts into a non-negligibleambient pressure. Examples of such rocket engines are large liquidpropellant sea-level rocket engines for boosters and core stages thatare ignited at sea-level and upper stage rocket engines ignited duringstage separation.

Dynamic loads are due to the instationary nature of the thrust chamberflow during start and stop transients and during steady-state operationwith separated flow in the nozzle. The rules for and effects of suchflow separation have been studied and presented at the 30thAIAA/ASME/SAE/ASEE Joint Propulsion Conference, June 1994, Indianapolis,Ind., USA in a paper “Aero-elastic Analysis of Side Load in SupersonicNozzles with Separated Flow”, Volvo Aero Corporation, to which may bereferred. The dynamic loads are generally of such a magnitude, e.g. ofthe order of 50-100 kN, that they present life-limiting constraints forthe design of thrust chamber components. These constraints result inhigher weight for thrust chamber structural elements. Furthermore thelargest possible area ratio that can be used on the nozzle extension islimited by the requirement of attached flow during steady-stateoperation.

The final consequences of the dynamic loads are constraints on theoverall performance-to-weight ratio of the thrust chambers and asubsequent limitation of the amount of pay-load that can be deliveredinto orbit by the rocket launcher.

For eliminating the drawbacks of prior nozzles a great number oftechniques have been suggested which all, however, have turned out tohave themselves significant drawbacks in various respects. The maindifficulties refer to the function, performance, cooling andreliability.

Thus traditional bellshaped nozzles give a limited function andsubstantial start and stop transient loads. A dual bell nozzle alsosuffers from severe transient dynamic loads. External expansion nozzleshave been suggested but not been sufficiently tested. A bell nozzleequipped with trip rings reduces dynamic loads but with too large aperformance loss. Said nozzles also suffer from difficult coolingproblems. Finally, extendible nozzles and ventilated nozzles have beensuggested but both require mechanisms with functions that are notpossible to verify prior to flight.

The main object of the present invention now is to suggest a rocketengine nozzle structure which provides for an advantageous flow controlwithin the diverging outlet portion of the nozzle which makes itpossible to reduce the weight of the rocket engine nozzle and to gainincreased performance.

According to the invention this is achieved by a nozzle which issubstantially distinguished in that, circumferentially, said outletportion in axial section has a radius, the length of which varies.Preferably, the radius length varies periodically and most preferablyperiodically so as to create a polygonal circumferential shape of thenozzle.

According to the invention, the separated flow can be controlledsatisfactorily by this very limited non-axisymmetric modification of thenozzle wall contour. This modification exhibits no significant negativeeffects on performance, reliability, cooling and manufacturing aspectsof the nozzle outlet portion.

The invention thus provides for the design of a sea-level rocket enginenozzle with significantly higher vacuum performance through a largernozzle area ratio and reduced weight.

The invention will be further described below with reference to theaccompanying drawing, in which

FIG. 1 illustrates in a perspective view the diverging outlet portion ofa nozzle according to the invention,

FIG. 2 illustrates a similar but opened or broken-up perspective view ofa similar nozzle, and

FIG. 3 is an axial section of a rocket engine with an outlet portionaccording to the invention.

With reference to FIG. 3 the invention will be described applied to aknown sea-level rocket engine. From the minimum nozzle area A in whichthe exhaust gases have sonic speed, the diverging or outlet portion ofthe nozzle can be considered to extend from the level B having an inletarea ratio of 5. Under many conditions such as in the moments ofstarting and stopping and when atmospheric pressure resides, the flowwill separate from the inner wall of said diverging portion of thenozzle. This phenomenon occurs spontaneously and randomly and causesinstationary gas loads in transversal direction. Said phenomenon willimply restrictions as to the expansion ratio in the outlet portion andthe length of the latter along which the requirement of attached flowduring steady state operation can be fulfilled. According to theinvention, a new nozzle shape is suggested which will provide for aneffective control of the flow separation in a reliable way such that theexpansion ratio might be increased and thus the power of the rocketengine without the flow separation phenomenon causing unacceptableside-loads.

The present invention, leaving the prior axisymmetric shape of thenozzle wall and having the radius varying in length circumferentially,and preferably to form a polygonal contour, will drastically improve thebehaviour of the movements of the flow separation line and reduce theside-loads and hence allow a higher expansion ratio.

In FIG. 1 has been illustrated only the diverging or outlet portion ofthe inventive nozzle. In said Figure, the axisymmetric contour of thenozzle portion has been modified from the prior pure circular shapeindicated by e.g. lines D, by letting, in accordance with the invention,the radius length vary periodically and preferably so as to form apolygon with e.g. eight sides S1, S2 . . . The polygon beingcircumscribed the circular line D, an area augmentation of maximum e.g.6% can be achieved.

In a preferred embodiment, the contour at the inlet of said divergingnozzle portion, i.e. at the level B, see FIG. 3, will be circular, thuswith constant length of the radius, and the radius length then startingprogressively to vary with increasing distance axially from the level Band to reach its greatest variations close to the end of the divergingnozzle portion.

In FIG. 2 the perspective view of the nozzle has been broken in order toshow the flow separation lines F. From said Figure it is evident thatthe flow separation lines will form curved portions extending from eachcorner of the polygonal contour and being curved downstream to about thecenter line of each polygon side. This implies that the flow separationcan be effectively controlled and the detrimental side-loads avoidedalmost entirely. The inventive polygonal contour will thus induce anon-axisymmetric pressure distribution of the inner wall surface of thenozzle outlet portion. In FIG. 2, the lines of flow separation have beenillustrated for a set of chamber pressures during a start transient.Like the embodiment shown in FIG. 1, also in this embodiment thepolygonal shape has eight (8) sides and provides an area augmentation by6% over the circular area circumscribed by the polygon.

The number of sides in the polygon can be varied from 5 to 15, 8, 10 and12 having been subject to tests. In analyses no affects on theperformance data have been observed.

As an example, the measures of a prior nozzle might be exemplified asfollows, viz.:

Length 1, 8 m Inlet diameter 0, 6 m Exit diameter 1, 8 m Inlet arearatio 5 Exit area ratio 45.

Contrary to this, the measures of a nozzle according to the inventionmight be exemplified as follows, viz.:

Length section A to exit 3, 3 m Inlet diameter 0, 6 m Exit diameter 2, 7m Inlet area ratio 5 Exit area ratio 100

As can be seen from the above dimensional example the exit area ratiocan be increased to 100 and thus provides for a higher expansion ratioand hence improved performance of the nozzle since the separation linesnow can be effectively controlled and moved further downstream of theoutlet portion of the nozzle to about a level C as illustrated in FIG.3. This also allows an increase in length of the nozzle outlet portionand improved flow conditions during steady state operation.

What is claimed is:
 1. A rocket engine system comprising: a singlerocket engine nozzle, extending from a rocket engine to an ultimatedownstream terminus of a rocket vehicle, with an outlet portion havingan axis and a curved profile in axial section, which outlet portion intransverse section has a radius with a length that varies periodically,wherein said length variation increases distally along the axis to amaximum such that a polygonal circumferential shape is obtained. 2.Rocket engine nozzle according to claim 1, wherein said polygonal shapehas between 5 and 15 sides.
 3. Rocket engine nozzle according to claim1, wherein the periodical variation of the length of the radius is smallor almost none closest to the smallest cross-sectional area of thenozzle and then progressively increases to the maximum desired finalvalue towards the mouth of said outlet portion.